The present invention relates to a hydraulically actuated aircraft engine control system, more particularly such a system interconnected with the aircraft hydraulic control system.
Present day aircraft gas turbine engines are highly sophisticated and, in order to improve their performance in regard to both thrust and fuel consumption, they include systems for controlling variable position compressor stators, variable geometry of the intake casing and devices to control and minimize the clearances between the stationary and rotating portions of their turbines and compressors.
Typically, these items are controlled by actuators in which the motive power is generated by pressurized fuel. The supply of pressurized fuel is delivered to the engine by one or more fuel pumps such that a portion of the fuel is supplied to and burned in the combustion chamber of the engine, while another portion of the pressurized fuel is directed to the control actuators and servo circuits for the variable engine structures. Unused fuel is directed by a control valve to the upstream side of a low pressure fuel pump. Any malfunction of the servo circuit of the gas turbine engine will cause engine stoppage, thereby making it impossible to drive the actuators for the variable engine structures and the rotor/stator control actuators.
Modern aircraft typically have at least one hydraulic control circuit for controlling the aircraft control surfaces, the operation of the landing gear, etc. with an independent pressurized fluid source which delivers a pressurized hydraulic fluid to the control circuit to control these structures. The hydraulic fluid used in the aircraft hydraulic control circuit typically is incompressible and non-flammable and may be subjected to a substantially higher pressure than the fuel for the gas turbine engine.